The invention relates to the field of launch vehicles. More particularly the invention relates to launch vehicular configurations including X33 aeroshell, external tank, and bell nozzle rocket engines.
The recently conducted, space transportation architecture study reviewed a wide range of launch concepts all proposed as options to replace the space shuttle. An independent assessment of these concepts determined that none were able to satisfy the wide range of design requirements that are demanded of a system to replace the space shuttle. The majority of the proposed launch systems are also being designed to meet commercial space lift needs. This commercial market demands that an additional set of challenging requirements be met.
One of the most demanding of the system design requirements for a launch system that is capable of performing both Civil (NASA) and commercial missions is to limit the development cost for the complete system to, on the order of, $2 Billion dollars. Using present year dollars, this figure is less than 10% of the original development cost of the space shuttle system. A number of contributing factors limit the development cost that is allowable. There is a the high cost on return on investment to obtain development funds for high risk development. There is a long duration between initial investment and an operational system with associated revenues. There is a the relatively small addressable launch market once operations begin. And, there is the relatively unknown demand elasticity environment in the launch market.
In addition to requiring low development costs, any system that is to replace the space shuttle must also be extremely reliable. This reliability is required not only over the long haul but also from the onset of initial operations. In addition, the new system must meet new manned flight human rating safety standards established by NASA. One of these safety standards demands that the system provide the ability for safe crew escape capability throughout the entire flight trajectory. The new set of safety standards is considerably more stressing than those required of the current space shuttle system. In fact, no other transportation system requires full crew escape throughout the entire journey. Rather, current transportation systems, such as commercial aircraft, are designed to be inherently safe and reliable and also provide for graceful degradation rather than accept catastrophic failure modes in the design.
Independent of NASA safety needs, the numerous direct and indirect costs associated with a catastrophic failure event implies the need for an extremely reliable design that precludes, within practicality, catastrophic system failures. Examples of design features that help to prevent such catastrophic failures include: full engine out from liftoff, full vehicle abort capability throughout the entire mission, robust design and operating margins, and integrated vehicle health management. The integrated vehicle health management working in conjunction with the vehicle management system provides the capability to anticipate impending system failures and or react promptly to unanticipated failures and take appropriate action to mitigate the risk associated with such failure. Such a capability implies heritage of the major system components and subsystems so that nominal and off nominal operating conditions can be recognized and dealt with accordingly. The alternative to a large degree of design heritage is to conduct an extremely expensive flight test program in which the system is fully characterized before revenue generating flight begins. This later option is most likely prohibitive in the space launch industry due to the small launch market in which to amortize these costs. The integrated vehicle health management system also provides vehicle health information in order to enable rapid and low maintenance cost turnaround of the launch system in preparation for the next mission. This capability is needed to meet cost goals. A requirement to allow incremental flight test is also important and allows the launch system to be characterized prior to subjecting it to overly stressing conditions. This incremental approach allows operating changes or design fixes to be incorporated before design or fabrication problems are allowed to cause a serious failure.
Other performance related requirements relating to space lift capability and recurring cost must also be met. The space shuttle orbiter has an injected weight of approximately 170K pounds with the exact weight being orbiter specific, for an enclosed usable volume for propellants and payload of 11K cubic feet. Using a metric of the ratio of injected mass to enclosed volume, the shuttles orbiter metric is 15.5, and is unsuitable for housing both propellant and a sizable payload. The system must lift approximately 25K pounds of cargo and crew to the international space station orbit and also be able to return crew and cargo from the international space station safely to Earth. It is also highly desirable that the launch system be able to evolve to support future Mars exploration missions. This demands a much heavier lift capability to low earth orbit than required for the international space station alone. Furthermore, to justify even a modest investment in a new launch system, the recurring cost must be low on the order of providing commercial space lift prices of less than $2,500 per pound including fully amortized development and production costs as well as cost of financing and allowance for profit.
The ideal concept should also be capable of evolving into a third generation launch system. Such a system would be capable of providing extremely low launch costs, in the range of $100 to $500 per pound, with levels of operability and reliability comparable to those of aircraft. This set of almost mutually exclusive design requirements for low development and operation costs, extreme reliability and safety, heavy lift, evolutionary capability, and heritage with flight proven hardware could not be met by any of the launch systems recently proposed to replace the Shuttle. Consequently, the transportation architecture study concluded that because no near term launch system could satisfy all of these requirements, the U.S. should continue operation of the space shuttle well into the future.
The space shuttle orbiter employs a wing body configuration and was designed with structural and thermal protection materials available since the late 1970s. The consequence of orbiter design approach and then available technologies yielded a vehicle design with a poor figure of performance in terms of the ratio of internal volume compared to the overall dry mass of the vehicle. The internal volume is used principally to house either propellants or payload. The poor internal volume to mass ratio of the space shuttle orbiter dictated the need for a large quantity of propellant and thus liftoff mass to accelerate the space shuttle orbiter and the payload to orbital velocity. Ultimately, the poor liftoff performance led to a decision that a rocket booster stage would be required to provide both the majority of liftoff thrust and initial acceleration, or impulse, of the space shuttle system. Segmented solid rocket motor powered boosters were selected for this purpose. The solid rocket boosters were intended to provide lower development costs than would reusable liquid boosters. However, the solid rocket booster operations significantly increased both operating costs and catastrophic failure modes of the space shuttle system. The extensive solid rocket booster recovery and refurbishment activities required after each flight also contributed to the large operations cost increase associated with the solid rocket boosters.
The solid rocket booster also had other liabilities. Once lit, the solid rocket boosters could not be turned off until their solid propellants were fully consumed. This is a major contributor to the inability of the space shuttle orbiter and the crew to abort during the first several minutes of each flight when the solid rocket boosters are providing the majority of the thrust from the space shuttle system. Another key contributor to the inability of the space shuttle orbiter to escape either a failed solid rocket booster or external tank (ET) and safely return to the launch site is a consequence of the design of the space shuttle orbiter that was designed to not house any propellants to operate the primary rocket engines or the space shuttle main engines (SSMEs). As a consequence of this lack of available propellants in the space shuttle orbiter, the SSMEs could not provide thrust to separate and operate the orbiter once detached from the ET. Thus, no means was provided for the space shuttle orbiter and the crew to escape the remainder of the space shuttle system in an emergency.
During the design of the space shuttle much emphasis was placed on reducing the required surface area of the expensive and labor intensive to maintain thermal protection system. A decision to utilize an expendable ET to house the primary liquid oxygen (LOX) and liquid hydrogen (LH) propellants rather than house the primary propellants inside the space shuttle orbiter enabled a great reduction in the size of the thermal protection system that would otherwise be required. This design approach also contributed to reducing system mass. The lack of experience in reuse of the primary LOX and LH propellant tanks further dictated the need to house primary LOX and LH propellants in an ET. These design choices led to the development of a complex space shuttle system that is still very expensive to operate, can only sustain a flight rate of approximately two flights per year for each orbiter, and is relatively intolerant of many critical failures of key flight systems.
Several years ago, the venturestar vehicle was proposed as a replacement to the space shuttle. The venturestar was envisioned to be capable of achieving single stage to orbit in which the entire vehicle and its payload could be accelerated to orbit without discarding any stages along the way. It was envisioned that such a system could address many of the design challenges associated with the space shuttle. The venturestar design employed a volumetrically efficient shape that significantly increased the ratio of internal volume of the venturestar compared to the dry weight of the venturestar. As indicated above, this improved figure of performance over the space shuttle orbiter enabled a reduction in the propellant mass required for the venturestar and its payload of approximately 25K pounds to be delivered to international space station orbit. If successful, this reduction in propellant mass could be used to entirely eliminate any type of propulsion to augment the thrust or propellant mass of the venturestar.
The shape of the aeroshell selected for venturestar combined with reduced reentry crossrange requirements yielded a reduction in the heat load that the venturestar would experience during reentry into the atmosphere of earth. This reduction in heat load allowed a metallic based thermal protection system design to be used on the venturestar. The metallic based thermal protection system is envisioned to be more robust and less costly to maintain than the ceramic thermal protection system currently used on the space shuttle orbiter.
NASA recognized that even with the major design and technology advances incorporated into the venturestar, the technological risk of achieving single stage to orbit was still very high. As a consequence, NASA jointly funded with industry the development of a subscale prototype named X33. The X33, rather than being designed as a vehicle that is itself capable of carrying payload to orbit, is under development as a demonstrator of the technologies and performance levels that would be required to demonstrate that the larger venturestar vehicle would be capable of delivering significant payload to orbit.
The X33 was chosen to be photographically half scale compared to the venturestar. As a consequence of surface to volume ratio scaling, the smaller X33 yields a reduced ratio of interior volume to dry mass than does the larger venturestar. As a consequence, the X33 was anticipated to be capable of demonstrating acceleration to a velocity of approximately 15K feet per second but with little or no useful payload. Such a level of performance was considered adequate to demonstrate that the larger venturestar would be able to achieve orbital velocity of approximately 25K feet per second plus approximately 25K pounds of payload.
As the X33 underwent development some of the key design approaches and technologies employed did not live up to expectations and subsequently the dry mass of the X33 increased. Due to internal volume and liftoff thrust limitations, the increased dry mass of the X33 yielded a reduction of the ultimate velocity that could be achieved with the X33. This consequence lent further doubt as to whether or not the venturestar goal of single stage to orbit was achievable. As a consequence, excitement for the development of venturestar and the precursor X33 demonstrator has waned.
Aerospike engines were selected to power the X33 and also the larger venturestar. The unique external expansion aerospike nozzle was the key component of an aerospike engine system. The improved external expansion aerospike nozzle was envisioned to provide a small reduction in the amount of propellant that would be required to achieve orbital velocity compared to that using more traditional bell nozzles. However, the aerospike design introduced several key disadvantages into both the design of the X33 and of the venturestar. The aerospike nozzle design required that pressurized gases be provided to fill the base area of the X33 vehicle. This was necessary to reduce base area induced drag to an acceptable level. Unfortunately, this requirement eliminated use of higher performing engine cycles such as the staged combustion cycle that is currently employed in the SSME and full flow staged combustion cycle such as used in the integrated powerhead demonstration.
These high performance engine cycles do not produce a pressurized exhaust gas stream that would be required to fill the base area of the vehicles. Therefore, a lower performing gas generator combustion cycle was selected for use with the aerospike nozzle. The combined effect of the more efficient aerospike nozzle and the lower performing gas generator cycle does not seem to offer a clear performance advantage over a traditional bell nozzle engine employing a staged combustion cycle at the vehicle system level. Furthermore, studies suggested that if an improved nozzle that was specifically designed to work with a staged combustion cycle were employed, such as a dual bell nozzle, any modest advantage of the aerospike engine would be further mitigated. It is therefore not clear that the widely claimed performance advantage of the aerospike engine will be practically realized.
More importantly, the inventor determined that the aerospike engine design employed on the X33 had one very negative performance characteristic that was not obvious to those skilled in the art. For aerospike engines to operate efficiently they require a much larger expansion area than does an equivalent thrust staged combustion engine and nozzle combination. To provide adequate thrust to lift the X33 its aerospike engines require virtually the entire base area of the vehicle to be used as an expansion surface for the exhaust plume of the engines. Specifically, the ratio of thrust produced per unit area of vehicle base area is significantly less than that achieved with existing staged combustion engines such as the SSME. One consequence of this is that the thrust output of the X33 is much more limited that would be possible if staged combustion engines were employed. Specifically, the two X33 aerospike engines are limited to a combined thrust of approximately 400K pounds.
As a consequence of the limited thrust capability and limited internal volume capability, the X33 was not designed to, nor could achieve orbital capability. It was also not clear that the significantly larger and more expensive venturestar could achieve the goal of single stage to orbit and therefore the single stage to orbit venturestar may never be built. Finally, the space shuttle system of which the venturestar was to replace, continues to be costly to operate and also suffers many critical failure modes. These and other disadvantages are solved or reduced using the invention.
An object of the invention is to provide a launch vehicle having an efficient fuel usage for payload to orbit.
Another object of the invention is to provide a launch vehicle having a bell nozzle rocket engine coupled to existing X33 aeroshells for efficient delivery of a payload into orbit.
Yet another object of the invention is to provide a launch vehicle having an external tank coupled to existing X33 aeroshells for efficient delivery of a payload into orbit.
Still another object of the invention is to provide a launch vehicle having an external tank and bell nozzle rocket engines coupled to existing X33 aeroshells for efficient delivery of a payload into orbit and to provide flight reentry of an orbiter.
A further object of the invention is to provide a launch vehicle having an external tank and bell nozzle rocket engines coupled to existing X33 aeroshells for efficient delivery of a payload into orbit.
The present invention is directed to a class of launch vehicles characterized by one or more rocket stages having bell nozzle rocket engines for thrusting with the rocket stages using X33 aeroshell flight control surfaces, and characterized by an attached propellant feeding stage for the efficient delivery of a payload into space. The rocket stages can be in the preferred forms an orbiter, a booster, two boosters or a booster and an orbiter. The orbiter preferably includes a payload bay. The propellant feeding stage in the preferred forms can be an external tank (ET) or a core stage the later of which preferably includes two bell nozzle rocket engines and a payload bay. The bell nozzle engines are preferably space shuttle main engines (SSME), NK33 rocket engines or RS68 rocket engines. The use of these components provides a variety of launch systems having a wide variety of capabilities.
The first form of the invention is a single orbiter five rocket engine launch vehicle. The orbiter includes five SSMEs and is attached to an ET that is shorter in length and to accommodate modified load paths into the orbiter. The orbiter uses an X33 aeroshell to provide a large internal volume but having reduced dry weight. The launch vehicle mates an X33 aeroshell orbiter to a modified ET. The single orbiter five rocket engine launch vehicle and the ET house additional propellants that would be required to lift the X33 aeroshell orbiter as well as a useful payload to a desired orbital velocity. To provide adequate thrust to launch the launch vehicle, five SSMEs are used to offer higher performance measured in thrust per unit base area. The large space available on the base of the X33 aeroshell orbiter is sufficient to enable locating up to five SSME along the base of the X33. Thus, the thrust level of the X33 aeroshell orbiter is increased to over 2000K pounds. The liftoff mass is constrained so that with the loss of thrust of a single engine at liftoff adequate thrust would still be provided to adequately lift the launch vehicle from the launch pad and enable a safe return to launch site. A significant portion of the interior volume of the X33 aeroshell orbiter is reallocated to provide payload carrying volume rather than providing volume for propellants. As a consequence of the high performance of the X33 aeroshell orbiter when reoutfitted with SSME engines, the internal volume of the modified ET provides sufficient capacity to house the majority of the primary propellants to achieve orbital velocity.
In a second form, a three rocket engine orbiter with a four rocket engine booster are both attached to an external tank that communicates propellants to the rocket engines of the orbiter and booster. In this preferred form, the four rocket engine booster is used to augment both the thrust capability and the propellant load of the launch system thus significantly increasing the amount of payload that can be lifted. Virtually all of the available internal volume of the four rocket engine booster is used to house propellants that are used during the initial vehicle ascent. Once at a velocity of approximately mach three and exhausted of propellants, the four rocket engine booster glides back to the launch site. The bulk of the interior volume of the three rocket engine orbiter is dedicated volume for payload with a small portion allocated for propellants. During a nominal mission, these propellants are used to provide the final acceleration of the three rocket engine orbiter to orbital velocity. As with the first preferred form, when it is desirable to lift the ET to orbit this can be done with a corresponding reduction of payload delivered on that mission. In a failure scenario, the propellants contained in the three rocket engine orbiter permit the three rocket engine orbiter and the crew to safely escape a failed four rocket engine booster and or a failed ET during any portion of the mission. In this second preferred form, under nominal circumstances, all seven rocket engines contribute to liftoff thrust. This level of output thrust permits an increase in the length and associated propellant capacity of the ET compared to the ET used in the previous preferred embodiment and to approximately the same length as the unmodified space shuttle ET. This level of performance from the four rocket engine orbiter and three rocket engine booster results in a payload lift capacity of approximately 95K pounds to the international space station. In addition to this increased payload capacity, the vehicle can also be sized to provide a margin to allow for positive vehicle acceleration with a failed engine. The inherent level of thrust provided by any six out of the seven available engines enables a full engine out abort capability right from liftoff. In the highly remote case of multiple engine failures or of a failed ET, the ET can be prematurely jettisoned allowing escape and recovery of both the three rocket engine orbiter and the four rocket engine orbiter. This wide range of tolerance to otherwise catastrophic failures greatly enhances the safety and reliability of launch.
In a third form, a two-two rocket engine orbiter with a four engine booster is attached to an ET. Two of the rocket engines on the two-two rocket engine orbiter utilize the LOX and LH propellants and are preferably SSMEs. The remaining two engines of the two-two rocket engine orbiter are referred to as hydrocarbon propellant engines that utilize LOX and a hydrocarbon propellant, such as a jet propellant or a rocket propellant, and are preferably NK33 engines. The interior volume of the two-two engine orbiter is comprised principally of payload volume, LOX volume and hydrocarbon propellant volume. This third form provides an alternate main propulsion system with increased tolerance to failures of the LOX and LH propulsion system. Nominally the two-two rocket engine orbiter stages from the ET at a velocity of approximately 16K fps and the two-two orbiter continues to orbit using only the LOX and hydrocarbon propellants contained within the two-two orbiter and then using only the two hydrocarbon propellant engines. This third form also provides the flexibility to carry the ET to orbit. When the launch vehicle delivers the ET to orbit, the propellant of ET is injected into the two-two rocket engine orbiter and four rocket engine booster until reaching final orbit. Because the higher injected performing LOX and LH propellants of the ET accelerate more mass to orbit, an increase in launch mass is realized. However, this increase in injected mass is approximately offset by the need to also carry the ET to orbit. On missions in which it is desirable to bring the ET to orbit, the X33 orbiter compromises the ability to escape from the ET during the final portion of the ascent. The LOX hydrocarbon propellant engines on the two-two engine orbiter can also provide orbital maneuvering capability to eliminate a separate maneuvering subsystem.
In a fourth form, two X33 four engine boosters are attached to a core stage to provide ultra heavy lift with a full engine out capability. The two X33 boosters are used to augment the lift capability of the core stage that includes a payload bay. The core stage launch vehicle includes the two X33 boosters with four SSMEs each and the core stage that houses ET type propellants tanks with two high thrust output expendable rocket engines such as low cost, high thrust RS68 rocket engines. The core stage also has a forward payload fairing that houses the payload in the payload bay. In this fourth form, both of the boosters have internal volumes dedicated to housing propellants that can be communicated with the core stage.
The preferred forms of the invention offer multiple X33 derived launch vehicles. Each of the X33 vehicles is retrofitted with multiple flight proven SSMEs that provide substantially more thrust. The increased lift capability provided by the higher thrust SSMEs allow the modified X33s to lift the existing, but slightly modified ET. In the flight arrangement, the X33 stages are mounted to the ET or core stage. The new launch vehicles integrate modified X33 hardware, SSMEs, NK33s, and a modified ET or core stage for reduced development costs. The modified X33 orbiter is projected to have an injected weight of approximately 95K pounds and provides an internal volume of approximately 12K cubic feet. This metric is calculated after increasing the X33 injected weight by 17K pounds to account for the design modifications to the X33 orbiter. The ratio of injected mass to enclosed volume is 7.75 for the X33 orbiters.
Excluding the benefit of the additional volume for propellants in the X33 orbiter, the reduced injected mass allows 75K pounds additional payload to be delivered to orbit. The propellants onboard the X33 orbiter are sufficient to impart the final delivery delta velocity as much as 10K feet per second. This allows the ET to separate and reenter well before having to be brought to orbit. Because the ET and residual propellant mass is on the order of 70K pounds, eliminating the need to carry the ET all the way to orbit increases payload by approximately 35K pounds. These new configurations offer as much as 95K pounds of payload that can be delivered to international space station orbit using the three rocket engine orbiter and four rocket engine booster design.
By using well proven and better performing SSMES, the thrust of the X33 derived vehicles is greatly increased. The three, four and five SSMEs and two-two engine configuration can be incorporated into the existing base area of the X33 while providing sufficient access to allow accommodation for engine servicing. The approximately 14xe2x80x2 foot high and 36xe2x80x2 foot wide base on the X33 vehicle provide sufficient space for up to five SSMEs. With the 7.5 foot diameter nozzles of the SSMEs, using a staggered configuration, space exists for both an upper and lower row of SSME engines on the base of the X33 booster and orbiter. Using five engines on each booster and the two engine core stage, a combined thrust of over six million pounds can be achieved. With a fully loaded core stage and two fully loaded X33 derived vehicles, this configuration can deliver a substantial mission payload of more than 200K pounds. An orbiter and a booster configuration both with three engines can be used to provide a reduced payload design of approximately 50K pounds. The invention is described in reference to four preferred forms but many variants are possible. These and other advantages will become more apparent from the following detailed description of the preferred embodiment.